This invention relates to gas turbine engines, and more specifically to the reduction of thermal loading on the forward outer seal and turbine disk of gas turbine aircraft engines. The invention is disclosed and explained in this application with specific reference to a first stage high pressure turbine (“HPT”) disk of gas turbine aircraft engines. Standard HPT rotor architecture of some gas turbine engines includes a forward outer seal that is carried entirely by the first stage HPT disk. This architecture has advantages, but one disadvantage is that the disk and forward outer seal are closely mechanically coupled but have different thermal response rates. During takeoff, the forward outer seal, which is very thin with a low mass, responds very quickly to an increase in air temperature, while the first stage disk, being much more massive, responds very slowly. Thus, the forward outer seal attempts to grow out at a relatively high rate in correlation with the increase in temperature, while the first stage disk grows much more slowly. The forward outer seal is therefore restrained from growing at a rate correlated to its increase in temperature, resulting in a thermal mismatch and large, thermally-induced loads on both the forward outer seal and the disk. This can result in low cycle fatigue life predictions for the forward outer seal.
Current practices to reduce or compensate for the thermal mismatch include adjusting the interface gaps between the forward outer seal and first stage disk and otherwise optimizing mechanical features of the forward outer seal in order to obtain acceptable LCF life. However, these practices have reached or neared their practical limits, so that further improvements using these techniques no longer appear likely. This application discloses a way of reducing basic thermal loading on the forward outer seal and first stage disk as a means of extending the LCF life to an acceptable extent.